The present invention relates to the specific field of turbomachines and more particularly it relates to the problem posed by mounting a combustion chamber made of a ceramic matrix composite (CMC) type material in the metal casing of a turbomachine.
Conventionally, in a turbojet or a turboprop, the high pressure turbine (HPT) and in particular its inlet nozzle, the combustion chamber, and the casing (or xe2x80x9cshellxe2x80x9d) of said chamber are all made of the same material, generally a metal. However, under certain particular conditions of use implementing very high combustion temperatures, using a metal chamber turns out to be completely unsuitable from a thermal point of view and it is necessary to make use of a chamber based on high temperature composite materials of the CMC type. Unfortunately, the difficulties of working such materials and their raw material costs mean that use thereof is generally restricted to the combustion chamber itself, while the high pressure turbine inlet nozzle and the casing continue to be made more conventionally out of metal materials. Unfortunately, metal materials and composite materials have coefficients of thermal expansion that are very different. This gives rise to particularly severe problems in making connections between the casing and the combustion chamber and at the interface with the nozzle at the inlet to the high pressure turbine.
The present invention mitigates those drawbacks by proposing a mounting for the combustion chamber in the casing that has the ability to absorb the displacements induced by the different coefficients of expansion of these parts. An object of the invention is also to propose a mount that enables manufacture of the combustion chamber to be simplified.
These objects are achieved by a turbomachine comprising an annular shell of metal material containing in a gas flow direction F: a fuel injection assembly; an annular combustion chamber of composite material having a longitudinal axis; and an annular nozzle of metal material having fixed blades and forming the inlet stage of a high pressure turbine; wherein said composite material combustion chamber is held in position in said annular metal shell by a plurality of flexible metal tongues regularly distributed around said combustion chamber, each of said tongues comprising three branches connected in a star configuration, the ends of two of the three branches being securely fixed to a downstream end of said composite material combustion chamber remote from said injection system via respective first and second fixing means, while the end of the third branch thereof is securely fixed to said annular metal shell by third fixing means, the flexibility of said fixing tongues making it possible at high temperatures for said composite material combustion chamber to expand freely in a radial direction relative to said annular metal shell.
With this particular structure for the fixed connection, the various kinds of wear due to contact corrosion in prior art systems can be avoided, and the presence of the elastic tongues replacing traditional flanges gives rise to an appreciable weight saving. In addition, because of their elasticity, these tongues can easily accommodate the differences of expansion that appear at high temperatures between parts made of metal and parts made of composite materials, while continuing to hold the combustion chamber properly and well centered inside the casing.
In a first embodiment, each of said first, second, and third fixing means is constituted by a plurality of bolts. In an alternative embodiment, only the third fixing means are constituted by a plurality of bolts, the first and second fixing means each preferably being constituted by a plurality of crimping elements.
Advantageously, the turbomachine of the invention further comprises a closure ring of ceramic composite material securely fixed to said downstream end of the combustion chamber, the ring being designed to form a bearing plane for a sealing gasket that provides sealing between said combustion chamber and said nozzle. Preferably, said closure ring is brazed to said downstream end of the combustion chamber. It may include a folded-back portion lying in line with the side wall of the combustion chamber.
In a first preferred variant embodiment, said bearing plane for the gasket lies in a plane perpendicular to said longitudinal axis of said combustion chamber.
In a second preferred variant embodiment, said bearing plane for the gasket lies in a plane parallel to said longitudinal axis of said combustion chamber.
In both these two variant configurations, the gasket is preferably of the omega type.
In a third preferred variant embodiment, said gasket is of the omega type. In this configuration, the gasket is preferably of the xe2x80x9cspring-bladexe2x80x9d type being held against said closure ring by means of a resilient element secured to said nozzle. Advantageously, the gasket can have a plurality of calibrated leakage orifices.